Supersonic cascades



W. DETTMERING May 6, 1969 SUPERSONIG CASCADES Sheet Filed July 20, 1967RIGHT ANGLE THRUS T INVENTOR Wilhelm Dettmerz'r BY Y M, fl'mufl'ATTORNEYS y 1969 Q w. DETTMERING 3,442,441

SUPERSONIC CASCADES I lfiled July 20, 1967 Sheet 13 of 2 INVENTOR BY W 4144 fiM ATTORNEYS United States Patent O Int. Cl. F04d 29736; F01d 5/06US. Cl. 230132 6 Claims ABSTRACT OF THE DISCLOSURE The gases flowingthrough a turbine or a compressor are reduced from supersonic tosubsonic flow by being passed through a row of shock and deflectingblades, and then through a row of gas deceleration blades whichpartially lap the shock and deflecting blades.

Because the stator of a turbine engine as, for example, an axialcompressor is subject to supersonic gas velocities, it is possible toincrease considerably the specific work turnover and therebyconsiderably reducing the crosssectional areas and the number of stages,and thus the axial length and material required for the obtaining of thedesired increase in pressure.

While the supersonic gas flow can be obtained with relatively goodefficiencies in an acceleration cascade, high losses occur in the heavydeflecting and deceleration cascades, such as disclosed in GermanPatents Nos. 759,892, 866,793 and 760,327, in which the gas flow isreduced to subsonic velocities by means of one or more shocks whichpreclude or heavily limit an economic use.

Aside from the unavoidable shock losses, there occurs a separation ofthe gas flow on the blade contours which considerably impairs theefliciency of a cascade. These losses which are influenced and caused bythe boundary layer, by the interaction between shock and boundary layer,and the deflection and deceleration to subsonic flow are considerablylessened according to this invention. According to this invention, thecascade is composed of a row of shock blades and a row of decelerationblades which partially lap each other and the blades are positioned withrespect to one another, as well as the contours of the blades so thatseparation of the gas flow cannot occur.

Double rows of blades have been used for the subsonic velocities, asdisclosed in German Patents Nos. 390,486, 459,204, 573,799 and 760,327.However, such double rows cannot be used for supersonic gas velocitiessince the blade contours and the duct cross-sections do not meet thecharacteristics of supersonic flow.

The means by which the objects of the invention are obtained aredescribed more fully with reference to the accompanying schematicdrawings in which:

FIGURE 1 is a cross-sectional view through a cascade showing one form ofthe invention;

FIGURE 2 is a similar view of a modified form of the invention; and

FIGURE 3 is another similar view of a further modified form of theinvention.

The very thin leading edges 1 and the trailing edges 2 of the shockblades 9 create only weak head waves with subsequent diagonal thrustswhich produce no substantial deceleration and cause but little loss instatic pressure. The vertical shock of the inflowing gas C is stabilizedby means of throttling within the area A so that a uniform cross-sectionof the passage in area A is required and at most the passage ispermitted to expand only slightly. The stream lines are not curved oronly slightly so since the vertical shock disturbs the equilibrium ofthe centrifugal force and pressure gradient. The deflection isaccomplished after the shock, that is with subsonic velocitiessubstantially in the area B. The losses connected wtih this are smallsince the boundary layers are relatively thin and since a decelerationand subsequent detachment of the flow is avoided by means of theapproximate uniform cross-section. The deflection is accomplished mostappropriately as a potential vortex. After the deflection and for thepurpose of building up the pressure, the cross-section in areas C and Dis enlarged and therefore the flow decelerated. The deceleration blade10 extends into the deflection area B to form areas C and D and is acondition for the exact adherence to the cross-sections which aresuitable for the flow. The trailing edges 2 of the shock cascade bladesare relatively thin in order to achieve a uniform annulus diagram and tominimize the low-energy wake flow.

The leading edges 3 of the deceleration cascade blades 10 are alsofeather-edged in order to avoid localized supersonic velocities. Theyare exactly tailored to the blades of the shock cascade and form asmooth blending of the cross-section, at the points where they overlapfrom areas B to C and D. The contours 7 and 8 of the deceleration blades10 form, with the contour on the suction side 6 and on the pressure side5 of the adjacent shock blades 9, a small angle of expansion ofapproximately 1 to 2 in area D and up to 4 in area C. The expansion ofthe cross-section in area D is held small because of the danger ofdetachment due to the continuation of the deflection which is necessaryin this type of cascade construction for geometric reasons and with dueconsideration for pressure equilibrium. The curvature in area Bproduces, by means of centrifugal force, a pressure gradient which,after the separation of the streams of gas flow in area C, effects ahigher means pressure than in area D. So that no pressure diflferencecan prevail at the confluence point of the two partial streams passingthrough areas C and D which could lead to a disturbance which couldcause a loss, the area D continues and ends with an inverted curvaturewhich again produces a pressure gradient which reaches at edge 2 themean pressure of the parallel flow in area C which is also freed from apressure gradient so that a uniform pressure and approximate parallelflow prevails upon the surface of discontinuity. Therefore, a smallpressure gradient remains only in area D which, however, causes nomentionable disturbance. In order to hold the pressure difference of thetwo partial flows low, it is not possible to expand the cross-sectionalarea C at will. The deceleration mainly takes place in the area E whichis kept as short as possible. At this point, a relatively high angle ofexpansion can be provided since the boundary layers are thin. Theaccumulated boundary layers of shock blades 9 flow into the sound coreof the flow and do not cause any appreciable loss. The outflowing gasesC pass between the trailing edges 4.

The curvature proposed for the equilization of pressure on the surfaceof discontinuity in passage E can be eliminated if the geometry of thecascade is otherwise the same and where the construction for a smallerdeflection in the cascade does not result in a too high pressuredifference. In this case, the suction-surface contour 6 of the shockblades 9 would extend in a straight line as shown in FIGURE 2. Thelosses will remain low since a type of stagnation point flow is formedon the trailing edge 2 which guarantees a continuous transition.

-In the modification of FIGURE 2, the parallel direction of the twopartial flows on the surface of discontinuity from edges 2 iseliminated. However, the additional losses produced by this fact areneutralized by the low inner deflection in the potential vortex areawith the angle of inlet and outlet remaining the same. Thus, withidentical effective deflection, the danger of detachment is reduced andthe cascade shortened. The contour 7 of the deceleration blade is on oneside constructed as an almost straight line while the curvature requiredfor the thickness distribution on the contour 8 produces a furtherdeflection.

In the modification of FIGURE 3, the shock is separated from the maindeflection and the subsonic deceleration. The last two are both placedin the second subsequent deceleration cascade where the boundary layershave not yet formed any dangerous thickness. As far as the constructionof the cross-sections is concerned, the above-mentioned points arevalid.

The shock cascade blades 9 are very thin over their entire length sinceotherwise even with a small positive angle Ant of attack and/ortransonic inlet flow velocity the cascade can be blocked, since thesupersonic flow conditions are not maintained. Small angle [3 of theleading edges already cause a pre-compression which has little loss,whereby the efficiency of the cascade is enhanced rather than when thevelocity change would be accomplished only in a vertical shock. However,because a detached head shock can occur which is connected with highlosses, the angle ,8 cannot be made much more than 6. The actualvertical shock is stabilized in the vicinity of the narrowestcross-section in order to obtain minimum static pressures. A lowdeflection of the flow by is also accomplished in the shock cascade bydesigning one contuor side straight, as has already been described forthe cascade in FIGURE 2. An insignificant deceleration in area B causedby the small angle 7 of expansion does not lead to any detachment.

The blades of the shock cascade 9 have in correspondence with their taskonly a small length so that, despite the shock, no dangerous thickboundary layers are produced. The boundary layer flows into the soundstream between the deceleration blades 10. The deceleration blades donot need to extend so far into being lapped with the shock cascadesince, with the small angle 7 of the trailing edge 2 of the shockblades, no greater expansion of the cross-section occurs which must beavoided.

Sharp leading edges 3 and small leading edge angles are needed becauseof the desired constant transition of the cross-section withoutrestriction. The main deflection takes place in the area E having analmost constant cross-section and with subsonic flow. The furtherdeceleration and pressude build-up occurs in the difiusion part D of thedeceleration cascade.

In the examples given, the flow has always been discharged in an axialdirection. This is not entirely necessary inasmuch as constructions arepossible which have a flow component in the direction of thecircumference. However, in order to fit to a subsequent stage or forachieving a discharge loss which is as small as possible, an axialdischarge is most appropriate in the majority of cases.

Even though the example given relate to a compressor, this inventionalso applies to a turbine subject to super sonic flow velocities andindeed also as the last stage for the purpose of retarding the highvelocity discharge and for deflecting the flow in the axial direction soas to considerably reduce the discharge losses.

This invention also applies to stationary as well as rotating cascadesfor considerably diminishing flow losses and realizing a highlyeffective specific work turnover in which the additional losses producedby the separation of the flow, by thick boundary layers and by theinteraction between shock and boundary layer are at least held to aminimum or avoided. Furthermore, it is possible to obtain largerdeflections and decelerations.

Having now described the means by which the objects of the invention areobtained,

I claim:

1. A deceleration lattice for supersonic fluid inflow in the stators androtors of axial flow compressors and turbines comprising two fixedstaggered blade rows arranged in series and with one row lapped with andstaggered downstream of the other, said staggered blade rows forming anupstream shock inducing cascade row and a downstream subsonic flowproducing cascade row, the blades in the upstream shock cascade rowhaving sharp leading and trailing edges and being spaced from each otherto form a fluid flow passage starting with a straight entry zone portionfollowed by a stabilizing slight narrowing of said zone portion andfurther followed down- I stream by a short curved deflection zoneportion having a nearly constant deflecting cross-sectional area fordeflecting the fluid flow toward said subsonic cascade row and saiddeflection zone portion ending at the leading edge of the blades in saidsubsonic cascade row, the blades in said subsonic cascade row beingspaced from each other and overlapping the trailing portions of theupstream shock inducing cascade to form a fluid velocity decelerationcross-sectional area larger than said deflecting cross-sectional area,and sharp leading and trailing edges on said subsonic blades for givingan almost constant crosssectional distribution of the flow passage.

2. A lattice as in claim 1, each shock cascade blade being concave onthe trailing edge portion of its suction surface, and each subsoniccascade blade being convex on its pressure surface for causing adeflection of fluid flow in a direction opposite to that of the maindeflection of the fluid flow through the entire lattice.

3. A lattice as in claim 2, said passage having a widenedcross-sectional area in the portion lapped by the blades for obtainingan initial deceleration of the fluid flow.

4. A lattice as in claim 1, each shock cascade blade being rectilinearon its suction surface portion lapped by the subsonic cascade blade, andthe subsonic cascade blade being rectilinear on its pressure surfaceportion lapped by the shock cascade blade.

5. A lattice as in claim 1, said shock cascade blades together with saidsubsonic cascade blades forming a blade set with each shock cascadeblade suction surface terminating with approximately the outlet angle ofthe entire blade set and said shock cascade blade having a thickness forforming a different angle on its pressure surface, said ditference beingcompensated for the fluid flow deflection into said subsonic cascaderow.

6. A lattice as in claim 1, each shock cascade blade having arectilinear pressure surface adjacent its leading edge and forming asmall angle with its pressure surface for producing an oblique shockstarting at said leading edge in addition to the normal shock with onlya weak deflection due to the finite thickness of the blade, and with themain deflection occurring in said subsonic cascade row.

References Cited UNITED STATES PATENTS 1,771,711 7/1930 Hahn 103972,839,239 6/1958 Stalker 230- 2,974,927 3/1961 Johnson 230-120 3,156,40711/ 1964 Bourguard 230120 3,356,289 12/1967 Plotkowiak 2301 14 HENRY FRADUAZO, Primary Examiner.

U.S C1. X.R. 230-120; 253-78

